Gas turbine blades have historically used compressor bleed air as the cooling medium to obtain acceptable service temperatures. Cooling passages associated with this design technology are typically serpentine arrangements along the mean camber line of the blades. The camber line is the locus of points between the low pressure and high pressure sides of the airfoil. Adjacent radial passages are connected alternately at the top and bottom by 180 degree return U-bends to form either a single continuous passage, or independent serpentine passages, with the cooling air exiting into the gas path by one or a combination of the following schemes (a) leading edge holes, (b) hole exits along the trailing edge, (c) hole exits on the high pressure side and low pressure sides of the blade airfoil, and (d) tip, cap holes.
Each radial passage typically cools both the high pressure and low pressure sides of the blade airfoil. The specific geometry of each radial cooling passage is designed to balance the conflicting demands for low pressure drop and high heat transfer rate. Schemes used in the state of the art to enhance heat transfer rate include raised rib turbulence promoters (also known as trip strips or turbulators), passage crossover impingement, the use of impingement inserts, and the use of banks or rows of pins. These schemes increase the local turbulence in the flow and thus raise the rate of heat transfer. The effectiveness of open circuit air cooling is further improved by the coverage of the blade airfoil by an insulating film of air bled through openings in the airfoil surface. The disadvantage of using compressor bleed flow, however, is that it is inherently parasitic. In other words, turbine component cooling is achieved at the expense of gas turbine thermodynamic efficiency. Cooling schemes involving high pressure and high density fluids, such as steam, on the other hand, have not yet been employed for blade cooling or reduced to practice in commercially available gas turbines.